The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Among the engine components, relatively high temperatures are observed in the combustor section such that cooling airflow is provided to meet desired service life requirements. The combustor section typically includes a combustion chamber formed by an inner and outer wall assembly. Each wall assembly includes a support shell lined with heat shields often referred to as liner panels.
In typical combustor chamber designs, the liner panels have a hot side exposed to the gas path. The opposite, or cold side, has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contact the inner surface of the liner shells. Testing has shown that the traditional cooling patterns may not be sufficient to provide effective thermal protection to all areas of the liner panel array with the ongoing lower emissions requirements and higher combustor operational temperatures.